The present invention relates generally to the field of propellants suitable for use in hybrid rockets, and more particularly to propellants and a method of selecting propellants that exhibit high regression rates.
Two basic types of chemical rocket propulsion systems are widely used in the rocket industry; namely, liquid systems and solid propellant systems. In a liquid system, liquid oxidizer and liquid fuel are fed at high pressure to a combustion chamber where they mix and react producing high temperature, high pressure gases which exhaust through a converging-diverging nozzle producing thrust. The mixing of reactants requires a high performance pressurization system for the fuel and oxidizer which must often operate in a cryogenic environment at extreme pressures and mass flow rates. The injection system and combustion chamber require exotic materials, complex systems for cooling, and very high precision manufacturing techniques. All of these factors contribute to a high cost.
Solid propellant systems do not require the complex and expensive machinery of liquid systems. Nevertheless, solid systems are complicated, and are subject to the difficulties of producing crack-free, repeatable, fuel grains, and by the need to transport and handle explosive materials. In a manufacturing process that requires extreme safety precautions, solid fuel and oxidizer are intimately mixed and allowed to cure inside the rocket case producing an explosive fuel with roughly the consistency of plastic or hard rubber. Fuel grains which contain cracks present a risk of explosive failure of the vehicle and must be rejected, driving up the cost of manufacture. Upon ignition the solid fuel burns uninterrupted until all the fuel is exhausted.
An alternative chemical rocket which has been known since the 1930""s is the hybrid propulsion system. In the hybrid design one propellant is stored in the solid phase while the other is stored in the liquid phase. Thus the hybrid lies somewhere between the two basic chemical rocket designs just described. In most hybrid propulsion applications, the solid is the fuel and the liquid is the oxidizer. Reverse hybrids with the fuel in the liquid phase and oxidizer in the solid phase are also feasible and the present invention described below can be applied equally well to both types of hybrid systems.
A large variety of fuels, including trash and wood, have been considered for hybrid rockets but the most conventional fuel materials are polymers such as Plexiglas (polymethyl methacrylate) (PMMA), high density polyethylene (HDPE), hydroxyl terminated polybutadiene (HTPB), and the like. Typical oxidizers that are frequently used in hybrid rockets are liquid oxygen, hydrogen peroxide, nitrogen tetroxide, nitrous oxide and occasionally fluorine. With respect to the last point, the fuel combinations used for hybrids are similar in their chemical properties and energy densities to the fuels used in hydrocarbon fueled liquid rocket systems. Thus, in terms of exhaust velocity and specific impulse, the hybrid system is a closer relative to a liquid system than to a solid system. Solid rockets tend to use lower energy oxidizers and consequently they produce lower specific impulse.
In addition to having a higher specific impulse, some of the advantages of the hybrid rocket over the solid fuel rocket are:
The hybrid allows for thrust termination, restart and throttling capabilities,
The hybrid design lends itself to safe manufacturing, transportation and operation.
Hybrid motors are inherently immune to explosion,
The safety and simplicity of the hybrid leads to lower development costs for new systems and very likely lower operational costs,
The combustion products are generally very benign producing lower environmental impact.
The main advantages of the hybrid over the liquid rocket include:
Lower development and operating costs (life cycle costs),
Lower fire and explosion hazards,
Less complex design and potentially higher reliability.
The hybrid allows the addition of energetic solid components, such as aluminum or beryllium to the fuel.
A schematic of a typical hybrid propulsion system 10 with a pressurized oxidizer feed system is shown in FIG. 1. The feed system is comprised of a pressurization tank 12 that holds an inert gas at high pressure (such as Helium, Argon or Nitrogen), a valve (not shown) to pressurize the oxidizer tank 14, a main valve 16 to turn on the flow of oxidizer and an injection system 18. Alternatively, the gas pressurization system can be replaced with a turbopump. The other major components are the combustion chamber 20 which contains the solid fuel 22 and the nozzle assembly 24.
A sketch of the flame configuration in a single port hybrid rocket combustion chamber 30 is shown in FIG. 2. The single port combustion chamber 30 generally includes a pre-combustion chamber region 31 at the front end, a post-combustion chamber region 32 at the opposite end, and an elongated single port 33 extending between the ends. The oxidizer in the liquid phase is injected into the combustion chamber at pre-combustion chamber region 31. The injected oxidizer is gasified and flows axially along the port 33, forming a boundary layer 34 over the solid fuel 22. The boundary layer 34 is usually turbulent in nature over a large portion of the length of the port. Within the boundary layer 34 there exists a turbulent diffusion flame 36 which extends over the entire length of the fuel. The thickness of the flame is generally very small compared to the boundary layer thickness. The heat generated in the flame, which is located approximately 20-30% of the boundary layer thickness above the fuel surface, is transferred to the wall mainly by convection. Some heat is also transferred by radiation but this is usually relatively small compared to the convective heat transfer. In the conventional hybrid system depicted in FIG. 2, the wall heat flux evaporates the (generally polymeric) solid fuel and the fuel vapor is transported to the flame where it reacts with the oxidizer which is transported from the free stream by turbulent diffusion mechanisms. The unburned fuel that travels beneath the flame, the unburned oxidizer in the free stream, and the flame combustion products mix and further react in the post combustion chamber 32. The degree to which fuel and oxidizer are able to fully mix and react before exhausting through the nozzle 24 determines the combustion efficiency of the motor. The hot gases expand through a convergent-divergent nozzle 24 to deliver the required thrust.
It is important to note that, even though the geometry of a hybrid motor is similar to a solid motor, the combustion scheme is vastly different. In a solid rocket, the oxidizer and fuel are both stored in the solid phase next to each other for heterogeneous fuels and within the same fuel molecule for double base fuels. Consequently, the solid combustion takes place in a deflagration (premixed) flame that is closer to the surface compared to the hybrid diffusion flame. Also, in solid fuel systems there exists some heterogeneous phase (solid-solid, solid-gas) reactions at the surface. In short, the burning rate of a solid rocket is determined by the rate of homogeneous (gas phase) and heterogeneous chemical reactions.
In a hybrid system or motor, the burning rate is limited by the heat transfer from the relatively remote flame to the burning surface of the fuel. One of the physical phenomena that limits the burning rate in a hybrid motor is the so-called blocking effect that is caused by the high velocity injection of the vaporizing fuel into the gas stream. This difference in the combustion scheme of a hybrid motor significantly alters the burning rate characteristics compared to a solid rocket. Blocking can be explained as follows. Increasing the heat transfer to the fuel causes the evaporative mass transfer from the liquid-gas interface to increase. But the increased blowing from the surface reduces the temperature and velocity gradient at the surface thus reducing the convective heat transfer. The blowing also thickens the boundary layer and displaces the flame sheet further from the fuel surface leading to a further reduction in convective heat transfer. The position of the flame sheet and the shape of the thermal and velocity boundary layer is the result of a complex chemical and fluid mechanical balance between the oxidizer flow entering the port, the fuel flow produced by evaporation and the flow of combustion products. As a result, the burning rate is limited in a fundamental way which is difficult to overcome by either increasing heat transfer to the fuel or by a reduction in the fuel heat of gasification. Although radiative heat transfer from the flame does not suffer from the blocking effect it is usually small compared to the convective heat transfer. The upshot of all this is that the regression rate, defined as the recession speed of the solid surface of a conventional hybrid fuel is typically one-tenth or less than that of a solid rocket fuel.
For a given selection of fuels and oxidizer to fuel mass ratio, the thrust generated by a rocket is approximately proportional to the mass flow rate. Thus a given thrust requirement dictates the fuel mass flow rate that needs to be achieved. The fuel mass flow generation rate is a product of the fuel density times the regression rate, multiplied by the burning surface area. The fuel density is determined by the type of fuels. Generally, high thrust levels are required for a launch vehicle. For a hybrid rocket design based on a slow burning conventional fuel, high thrust can only be achieved by increasing the burning surface area. The high burning area requirements, and various other design constraints (such as the maximum grain length to port diameter ratio), leads to complicated multi-port configurations. One commonly used multi-port configuration is the wagon wheel geometry as shown in FIG. 3, and has been implemented in several hybrid motor designs.
The wagon wheel configuration and all other complicated multi-port designs have serious disadvantages. These disadvantages include:
the large sliver fractions, which may in practice leave significant amounts of fuel unburned;
fairly small volumetric loading of the fuel in the casing leading to decreased mass fractions;
grain integrity problems, especially towards the end of the burn when the web thickness between ports becomes vulnerable to structural disintegration;
difficult and expensive manufacturing of the fuel grain; and
requirement for a pre-combustion chamber or multiple injectors.
It is clear that all these factors seriously degrade the overall efficiency and cost of a multi-port hybrid launch vehicle.
The low regression rates and consequent multi-port design requirements make hybrids an unattractive option, even though they offer significant advantages over currently used liquid and solid systems. In order for the hybrid to find use as a practical design with a variety of applications, higher regression rates are required. Thus, so far many techniques have been suggested, or tried, to improve the regression rates of hybrids, however all of these techniques suffer important shortcomings. More specifically, one such prior art method uses fuels with low effective heat of gasification. This method yields only a small improvement since, as revealed in the classical hybrid theory (reference 1), the exponent of the heat of gasification is a small number (approximately 0.32). The weak dependency of the regression rate on the heat of gasification is due to the blocking effect described earlier. Other prior art techniques use insertion of screens in the port to increase the turbulence level, and thus the heat transfer rates. As with any method which requires that devices be placed in the gas flow path, this method complicates the design significantly and increases the likelihood of failure. In addition, this approach may lead to nonuniform burning along the port.
The addition of swirl to the incoming oxidizer flow to increase the effective mass flux and thus improve the heat transfer rate has also been reported (Reference 8). This method also complicates the hybrid design, especially for large scale motors, and requires heavy injectors or vanes.
Another prior art approach employs the addition of oxidizing agents or self decomposing materials in the hybrid fuel. This well known technique reverts to a quasi-solid design and eliminates the inherent safety characteristic of hybrid rockets.
The addition of metal additives has also been used. This is a common technique that improves the fuel mass burning rate. The improvement is small, however, and there are several shortcomings such as the increased vulnerability to instabilities due to the pressure dependent regression rate and increased environmental impact.
Yet another prior art technique focuses on increasing the roughness of the burning surface by adding dispersed phase particles in the fuel that would burn at a different rate compared to the matrix material (Reference 9). This technique can only give a limited improvement and large solid particles injected in the gas stream reduce the efficiency of the system. The manufacturing costs would also increase.
As just described, the prior art techniques are subject to significant limitations and disadvantages. Accordingly, it is highly desirable to provide a propellant and hybrid system which exhibits a high regression rate, without compromising safety or manufacturing cost.
[1] Marxman G. A., C. E. Wooldridge and R. J. Muzzy, xe2x80x9cFundamentals of Hybrid Boundary Layer Combustionxe2x80x9d, Progress in Astronautics and Aeronautics, Vol. 15, 1964 p 485.
[2] Karabeyoglu M. A., xe2x80x9cTransient Combustion in Hybrid Rocketsxe2x80x9d, Stanford University Ph.D. Thesis, August 1998.
[3] Gater R. A. and M. R. L""Ecuyer, xe2x80x9cA Fundamental Investigation of the Phenomena that Characterize Liquid Film Coolingxe2x80x9d, International Journal of Heat and Mass Transfer Vol. 13, pp 1925-1939, 1970.
[4] Ishii M. and M. A. Grolmes, xe2x80x9cInception Criteria for Droplet Entrainment in Two Phase Concurrent Film Flowxe2x80x9d, AICh Journal, vol. 21, no. 2, pp. 308-318, 1975.
[5] Nigmatulin R., B. Nigmatulin, Y A. Khodzaev and V. Kroshilin, xe2x80x9cEntrainment and Deposition Rates in a Dispersed-Film Flowxe2x80x9d, International Journal of Multiphase Flow Vol. 22, pp. 19-30, 1996.
[6] Bicerano, J, xe2x80x9cPrediction of Polymer Propertiesxe2x80x9d, Marcel Dekker Inc., 1996.
[7] Dauber, T. E., Danner, R. T., xe2x80x9cPhysical and Thermodynamic Properties of Pure Chemicals, Data Compilationxe2x80x9d, Taylor and Francis, 1997.
[8] W. H. Knuth, M. J. Chiaverini, D. J. Gramer and J. A. Saver, xe2x80x9cSolid-Fuel Regression Rate and Combustion Behavior of Vortex Hybrid Rocket Enginesxe2x80x9d, Thirty-fifth Joint Propulsion Conference and Exhibit, AIAA Paper No. 99-2318, 1999.
[9] D. B. Stickler, xe2x80x9cHeterogeneous Fuel for Hybrid Rocketxe2x80x9d, U.S. Pat. No. 5,529,648 issued Jun. 25, 1996.
[10] DeRose, M. E., K. L. Pfeil, P. G. Carric and C. W. Larson, xe2x80x9cTube Burner Studies of Cryogenic Solid Combustionxe2x80x9d, AIAA/SAE/ASME/ASEE Thirty-third Joint Propulsion Conference and Exhibit, AIAA Paper No. 97-3076, July 1997.
Accordingly, it is an object of the present invention to provide hybrid rocket propellants that exhibit a high regression rate, or more specifically, that will burn several times faster than conventional propellants at the same operating conditions of port mean mass flux and chamber pressure while retaining the basic advantages of hybrids; throttlability, safety and low cost. In addition to a high burning rate it is desired, but not required, that the propellant have the following characteristics:
self-decomposing materials are not involved;
the port design can be structurally simple;
the propellant is safe, easy to handle and easy to process;
the propellant burns smoothly; and
the burning rate is reasonably uniform along the axial and in the azimuthal directions in the port.
It is also an object of the present invention to provide a method of selecting, or identifying, such hybrid rocket propellants.
As described in the Background, in a hybrid rocket combustion chamber, liquid oxidizer is converted to gas and caused to flow over the solid fuel surface. In a reverse hybrid the oxidizer is the solid. Upon ignition, a flame sheet is formed above the solid surface and heat from the flame melts the solid causing a liquid layer to form. Evaporation from the liquid-gas interface produces a continuous flow of fuel gas which mixes with oxidizer at the flame sheet thus maintaining the combustion. At steady state, the regression rate of the melt surface and the liquid-gas interface are the same and the thickness of the liquid layer is constant. FIG. 4 shows typical steady state temperature and velocity profiles in a liquefying hybrid rocket.
The inventors have discovered, and according to the present invention, the liquid layer at the melt surface can be hydrodynamically unstable under the mass flux, pressure and temperature conditions which occur in a hybrid rocket combustion chamber. This shear-driven instability leads to wave formation on the liquid-gas interface and as the waves develop nonlinearly, the displaced liquid-gas interface exposed to the high speed flow of gas can breakup, leading to the formation of concentrated pockets of high density fuel and/or fuel droplets which are entrained into the gas stream. The mechanism of liquid layer instability and entrainment can substantially increase the rate of mass transfer from the fuel surface. This situation is illustrated schematically in FIG. 5.
We have developed a method for solid propellant selection or identification that takes the mechanism of liquid layer instability and entrainment into account. This method of the present invention has been used to identify high regression rate solid fuels and to predict their performance. It can be applied equally well to solid fuels or oxidizers which are collectively referred to as propellants. An important element of the process is a criterion that determines whether a given solid propellant is more or less likely to produce entrainment for a given set of combustion chamber conditions.
Accordingly, the present invention provides for a fuel composition suitable for use in hybrid rockets having a fuel component and an oxidizing component. One of the components flows past the other, and under the heat of combustion (heat transfer from the flame) one of the components forms an unstable melt layer with viscosity and surface tension such that droplets from the melt layer are entrained in the other component thereby increasing the burning rate. The present invention can also be used in formulating a fast burning fuel for solid fuel ramjet applications.
In another aspect of the present invention a propulsion system is provided. The propulsion system includes a vehicle structure, terminating in a nozzle and having a fuel component within the structure. One or more combustion chambers are formed within, or alternatively contain, the fuel component. Also provided is an oxidant vessel within the vehicle structure for flowing the oxidant in contact with the one or more combustion chambers to react with the fuel. The fuel is selected such that under the heat transfer from the flame, the fuel forms an unstable melt layer with viscosity and surface tension such that droplets of the melted fuel are entrained in the flowing oxidant thereby increasing the burning rate.
In yet another aspect of the present invention, a combustible hybrid fuel having a solid fuel component and a flowing oxidizer component flowing through one or more ports is provided. The solid fuel forms a liquid layer at the interface between the oxidizer and fuel, and the liquid layer exhibits entrainment of liquid droplets in the flowing oxidizer at an entrainment rate of             r      .        ent    ∝                                          (                                          C                f                            ⁢                              P                d                                      )                    α                ⁢                  h          β                                      µ          γ                ⁢                  σ          π                      ·  
In still another aspect of the present invention a method of selecting a propellant that exhibits a desirable regression rate during combustion within a port having a gas stream flowing through the port is provided. The method comprises the steps of:
for a given port mass flux, G=xcfx81gUg, where xcfx81g is the port average gas density and where Ug is the port average gas velocity; and
for a thickness h of a liquid layer formed on the surface of the fuel;
wherein the port mass flux value and the thickness satisfy the relationship of
G1.6h0.6xe2x89xa7aonset 
and where aonset is the entrainment onset parameter and is given by:       a    onset    =      1    .    05    xc3x97          10              -        2              ⁢          (                        ρ          g          1.3                          ρ          l          0.3                    )        ⁢          1                        (                                    C              fref                        ⁢                          C              B1                                )                0.8              ⁢          (              1                  µ          g                    )        ⁢    σ    ⁢          xe2x80x83        ⁢          µ      l      0.6      
and selecting the propellant such that aonset has a value that promotes entrainment of droplets from the liquid layer into the gas stream in the port.